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NASA Technical Memorandum 4538

Flight-Determined

Engine ExhaustCharacteristics of an

F404 Engine in an

F- 18 Airplane

Kimberly A. Ennix, Frank W. Burcham, Jr., and Lannie D. Webb

October 1993

(NASA-TM-4538) FLIGHT-DETERMINEO

ENGII_E EXHAUST CHARACTERISTICS OF

AN F40e ENGINE IN AN F-I_ AIRPLANE

(NASA) 15 p

N94-15141

Unclas

HI/O? 0189624

https://ntrs.nasa.gov/search.jsp?R=19940010668 2020-07-25T20:38:50+00:00Z

NASA Technical Memorandum 4538

Flight-Determined

Engine ExhaustCharacteristics of an

F404 Engine in an

F- 18 Airplane


Dryden Flight Research Facility

Edwards, California

National Aeronautics and

Space Administration

Office of Management

Scientific and Technical

Information Program

1993

FLIGHT-DETERMINED ENGINE EXHAUST CHARACTERISTICS OF AN

F404 ENGINE IN AN F-18 AIRPLANE

Kimberly A. Ennix*Frank W. Burcham, Jr.**

Lannie D. Webb***

NASA Dryden Flight Research FacilityEdwards, California

Abstract FG

HSCTPersonnel at the NASA Langley Research Center

(NASA-Langley) and the NASA Dryden Flight Research M9

Facility (NASA-Dryden) have recently completed a joint Mjetacoustic flight test program. Several types of aircraft withhigh nozzle pressure ratio engines were flown to satisfy a M,

twofold objective. First, assessments were made of sub- NPRsonic climb-to-cruise noise from flights conducted at vary-

ing altitudes in a Mach 0.30 to 0.90 range. Second, using P8

data from flights conducted at constant altitude in a Mach

0.30 to 0.95 range, engineers obtained a high-quality noise PLAdatabase. This database was desired to validate the Air-

craft Noise Prediction Program and other system noise Pamb

prediction codes. NASA-Dryden personnel analyzed the Ps9engine data from several aircraft that were flown in the

test program to determine the exhaust characteristics. The /8

analysis of the exhaust characteristics from the F-18 air-

craft will be reported in this paper. This paper presents anoverview of the flight test planning, instrumentation, test

procedures, data analysis, engine modeling codes, andresults.

Nomenclature

A8

AE8

AE9

ANOPP

exhaust nozzle physical area at the throat,in 2

exhaust nozzle effective throat area, in2

exhaust nozzle effective area at the exit

plane, in2

Aircraft Noise Prediction Program

*Aerospace Engineer.

**Chief, Propulsionand Pedormance Branch. AIAA AssociateFellow.

***Aermpaee Engineer. Member AIAA.

Copyright © 1993 by the American Institute of Aeronautics and Astro-

nautics, Inc. No copyright is asserted in the United States under "rifle 17,

U.S. Code. The U.S. Government has a royalty-free license to exerche

all rights under the copyright claimed herein for Govermnental proposes.

All other rights are reserved by the copyright owner.

V9

V.

W8

gross thrust, lb

high-speed civil transport

nozzle exit Mach number

fully expanded jet Mach number

free-stream Mach number

nozzle pressure ratio, P8/Pamb

total pressure at the exhaust nozzle throat,

psia

power lever angle, deg

ambient pressure, psia

smile pressure at the exit plane, psia

total temperature at the exhaust nozzle

throat, °R

nozzle exit velocity, ft/sec

fully expanded jet velocity, ft/sec

free-stream velocity, ft/sec

mass flow rate at the exhaust nozzle throat,

lb/sec

Introduction

Environmental issues are a significant concern con-

fronting the designers of future high-speed civil transport

(HSCT) airplanes. It has been determined that a substan-tial market will exist for HSCT aircraft if designers meet

key environmental issues, one of which is noise. The

HSCT aircraft must keep takeoff, climb-to-cruise, and

landing noise levels within the Federal Aviation Regula-

tion, part 36, Stage III community noise standards, l

The HSCT design concept will likely have supersonic

cruise speeds between Mach 2.00 and 2.50. Engines capa-

ble of efficient flights at speeds above Mach 2.00 will

likely have the thermodynamic cycle of a turbojet or a

very-low-bypass turbofan. 2 These engines have high

nozzle pressure ratios (NPRs) and jet velocities, whichraises concern not only for takeoff and landing noise, butalso for climb-to-cruise noise, extending from the airportfor a distance of up to 50 miles.

To determine the predicted noise of HSCT aircraft,acoustic codes such as ANOPP (Aircraft Noise PredictionProgram)3 are used. These codes were developed using

data acquired from engines with NPRs and flight speedslower than those planned for HSCT aircraft.

To better understand the acoustic characteristics of

engines representative of HSCT designs and to enhancecurrent noise prediction codes, personnel at NASA Lan-gley Research Center (NASA-Langley) and NASA Dry-den Flight Research Facility (NASA-Dryden) haveconducted joint flyover acoustic testing to acquire data.4

The test objective was first, to assess subsonic climlyto-cruise noise using aircraft with high NPR engines and sec-ond, to obtain an improved noise database to validateANOPP and other system noise prediction codes. TheNASA-Dryden personnel conducted the flyover tests anddetermined the engine exhaust characteristics. TheNASA-I_,angley personnel made the acoustic measure-ments, performed the correlations between the engineexhaust characteristics and acoustic data, and updated theANOPP code.

The flight study consisted of a series of flights overmicrophone arrays using several types of aircraft. In thesubsonic climb portion of the study, the flight matrix con-sisted of flyovers at intermediate power at altitudes from3800 to 32,000 ft and Mach numbers from 0.30 to 0.95.

For the ANOPP evaluation flyovers, the test points wereconducted at a constant altitude and lVL_h number. A

ground static acoustic test was also conducted to establishacoustic levels with no forward velocity.

For all of the tests, the measured engine data werecollected and later analyzed by an F404-GE-400 in-flightthrust code. The code predicted the engine exhaust char-acteristics of exhaust velocity and Mach number, whichcannot be directly obtained from the measured enginedata.

This paper describes the 1:-18 ahplane, the F404

engine, the flight test program, and the methods used tocalculate the engine exhaust properties. In addition, thepaper presents the exhaust velocity and Mach number datafor the climb-to-cruise, ANOPP validation, and groundtests.

Aircraft Description

F-18 Aircraft

Figure 1 shows an Fo18 aircraft (McDonnell DouglasCorp., St. Louis, Missouri and Northop Corp., Newbury

2

Park, California) during one of the test runs. This super-sonic, high-performance righted" has excellent transonicmaneuverability and is powered by two F404-GE-400(General Electric Co., Lynn, Massachusetts) afterbuming,turbofan engines. 5 Both engines are mounted close

together in the aft fuselage. The standard F-18 mainte-nance data recorder was used onboard the aircraft to

record a limited number of airplane and engine

parameters.

Engine Description

The F404-GE-400 is a two-spool, low-bypass, axial-flow turbofan with afterburner. The engine consists of athree-stage fan driven by a single-stage, low-pressare tur-bine and a seven-stage, high-pressure compressor drivenby a single-stage, high-pressure turbine. Variable geome-

try is incorporated into the fan and high-pressure compres-sor and the nozzle is a convergent-divergent nozzle. 6 It is

equipped with an engine control unit (ECU) where idlepower is 35° power lever angle (PLA) and intermediate(maximum nonafterburning) power is 102° PLA. It pro-daces NPRs similar to those expected of an HSCT engineapplication.

Test Procedures

Ground Track

The flight tests were flown over the Rogers Dry Lakewhich is adjacent to NASA-Dryden. Located at an ele-vation of 2300 ft, this dry lakebed provides a flat, interfer-ence-free area for acoustic testing. Figure 2 shows theapproximate location of the microphone array placedalong the "fly-by" line on the nortlw_t side of the lakebed.This area was ideal for tracking because of its close prox-imity to the NASA-Dryden radar site. Using the groundtrack and distance displayed in the control room, the pilotswere guided down the ground track and over the acousticarray. Flight conditions such as altitude or Mach numberneeded to be kept as constant as possible to get good quan-titative runs. There were 95 recorded F-18 flyovers.

Flight Procedures

The flight tests were conducted in two segments: sub-sonic climb to cruise and ANOPP validation. A single-exhaust jet was desired so the acoustics tests would haveone distinct noise source. For the twin-engine F-18 air-craft both engines were used to set up the initial conditionsbefore the beginning of a maneuver. On call from the con-trol room, the pilot stabilized the speed and altitude of the

aircraft. The left engine was throttled back to idle whilethe right test engine was operated at intermediate power.This procedure simulated the effect of a single engine. Forthe ANOPP validation segment the test engine was

operated at the power required to maintain level flight.The F-18 maintenance recorder was activated by the pilot

and operated for 35 see to record aircraft and engine dataduring the run. Table 1 shows the flight matrix for theclimb-to-cruise and ANOPP validation segments of the

flight test.

Climb-to-Cruise Test Matrix

The flight matrix for the climb-to-cruise segment con-

sisted of level flight accelerations at various Mach num-

bers and altitude to simulate points along a expected

HSCT climb profile. Altitudes varied from 3800 to 32,500

ft and speeds ranged from Mach 0.30 to 0.90. The aircraft

flew with the right test engine at the intermediate power

setting to maximize NPR.

ANOPP Validation Test Matrix

The ANOPP evaluation segment was flown at a con-

stant altitude of 3800 ft (1500 ft above the ground) with

speeds ranging from Mach 0.30 to 0.95. Power settings onthe test engine varied depending on what was required to

maintain constant flight speed or Mach number and

altitude.

Ground Test

In addition to the flight testing, static ground testswere conducted with the aircraft tied down on the thrust

stand pad at the US Air Force Flight Test Center atEdwards, California. The test matrix consisted of PLAs

from idle to intermediate power at 0.1 increments in

engine pressure ratio. The F-18 onboard recorder mainte-nance tape was run for 35 sec at each power setting to

record the engine data. Temperature, wind speed, andwind direction were also recorded. The tests were con-

ducted with the wind speeds below 5 kts.

Engine Exhaust Characteristics

Jet-mixing and shock cell noise are the two jprimarysources of noise for takeoffs and subsonic climbs." These

noise sources are primarily affected by the aircraft veloc-

ity, the exhaust exit Mach number and velocity, and theNPR. For acoustic analysis, engine exhaust characteris-tics are often defined at the nozzle exit and an assumed

fully expanded jet. Jet-mixing noise is a function of thedifference between the fully expanded jet velocity (V#t)and the free-sl_eam velocity (V**). Shock cell noise is a

function of the difference between the fully expanded jet

Mach number (Mjet) and the nozzle exit Mach number

(M9). As you approach the point where M 9 = Mja and V**

= Via" the shock cell noise and jet-mixing noise are dimin-ished. Nozzle exit velocity (V9) and M9 are based on the

aerothermodynamic characteristics of the flow at the noz-

zle exit plane (Fig. 3).

F404 In-Flight Thrust Code

Data obtained from the engine during the flight and

ground tests included compressor speed, compressor dis-

charge pressure, fan speed, fuel flow, inlet temperature,

turbine discharge temperatures, turbine discharge pressure,and nozzle area ratio. Measured engine data obtained

from the flight tests do not directly give the values of M 9,

V 9, M#t, and V#t needed for acoustic analysis with theANOPP prediction code. As a result, the measured engine

data must be input into the engine performance codes to

obtain the desired engine exhaust characteristics.

The F404-GE-400 in-flight-thrust performance code s

was developed by the General Electric Co. for the US

Navy. This code models the engine as a gas generator anduses the measured engine parameters as input. The perfor-

mance code calculates the following parameters through-

out the flight envelope: gross thrust (FG), Vg. Via, Mg,

M#t, NPR, exhaust nozzle effective exit area to effectivethroat area ratio (AEg/AE8), exhaust nozzle static exit

pressure to ambient pressure ratio (Ps91Pamb), exhaustnozzle throat total temperature (7'8), and exhaust nozzle

throat mass flow rate 0V8). The exhaust nozzle exit mass

flow rate(w9) andtotaltemperature(T9)areassumed to

be equal to W8 and I'8, respectively.

The following assumptions are used in the in-flight

thrust code. Steady one-dimensional isentropic flow isassumed between the throat and the nozzle exit. Based on

the resulting nozzle static exit pressure (Ps9), the flow

will be overexpanded (Ps9 < Pamb), fully expanded (Ps9

= Pamb), or underexpanded (Ps9 > Pamb). Mje t is based

on the point where the flow is fully expanded (Ps9 =Pamb) and it is a function of NPR. M 9 is a function of

nozzle area ratio. Once M 9 and Mje t are determined, V 9

and Via are then calculated. V 9 represents the actualexhaust exit velocity while Vje t represents the ideal fully

expanded jet exhaust velocity. If the actual exhaust veloc-

ity were fully expanded, V 9 would match V#t.

Results and Discussion

Climb-to-Cruise Test Results

Figure 4 shows the effect of Mach number on F404

engine exhaust characteristics for climb-to-cruise tests atintermediate power. Figure 4(a) shows the relationship

between V#t and V9 and the free-stream Mach number(M**). Each point on the curve represents a differentaltitude in the climb-to-cruise matrix. At the beginning of

the climb profile when the altitude is approximately 3800

ft and M,, = 0.30, the nozzle is overexpanded (V 9 > V#t).

The point where the data crosses, M** = 0.85 and V#t = V9.

indicates that the nozzle is fully expanded. When theclimb profile reaches an altitude of approximately32,300 ft and M** - 0.90, the nozzle is underexpanded

(V9 < l_et). Overall, V9 varies from V9 ,- 2750 ft/sec up toa maximum of V9 - 2800 ft/sec, and then drops to a value

of V9 = 2750 ft/sec, while Viet varies from 2300 to 2900ft/sec.

Figure 4(b) shows Mje: and M 9 as a function of M,,.Mje t and M 9 follow the same trends with Mach numberand altitudes as Via and V9. The values of M 9 varybetween 1.69 up to a maximum of M 9 _ 1.80, and then

drop to M9 = 1.70. The values of Mje t vary betweenMje: ,_ 1.35 up to a maximum of Mj¢t = 1.76. Above M,,

0.85, the difference between the two values is signifi-cantly reduced. Table 2 lists other parameters of interestfor the climb-to-cruise test. The maximum nozzle pres-sure ratio was 5.24.

ANOPP Validation Test Results

Figure 5 shows the effect of Mach number on F404engine exhaust characteristics for ANOPP validation testsat a level altitude of 3800 ft. Figure 5(a) shows the

exhaust velocities V9 and Vjet with respect to M**. Thepower setting (PLA) of the test engine was set at the level(shown in parentheses) necessary to maintain constantMach number in level flight while the other engineremained at idle. The power settings varied from partialpower (75 °) at the lower speeds, to intermediate power

(102 °) at the higher speeds. As in Fig. 4, both V9 and Vjetwere plotted as a function of M,,; however, no crossoveroccurred because at this low altitude the nozzle is overex-

panded for this M**range. The values of V9 varied from

V9 = 2550 ft/sec up to a maximum of V9 ,= 2900 ft/sec,

while Vjet varied between Via _ 1900 to 2650 ft/sec.

Figure 5(b) shows a plot of exhaust Mach numbers

M 9 and Mjet with respect to M,_ This set of data alsoshows a steady trend of increased M 9 and Mjet without anycrossover of the data. The two curves do converge towardeach other indicating that a fully expanded nozzle condi-tion may occur at a higher M**. The values of M9 were

M 9 = 1.70 to 1.80 while M_t varied from Mjet ,= 1.15 to1.60. Table 3 shows additional parameters of interest forthe ANOPP validation test.

Ground Test Results

For the ground static tests the effect of PLA on F404engine exhaust characteristics is shown in Fig. 6. Figure

6(a) shows V 9 and Viet plotted against PLA for theground tests. The values of V9 varied from V9 -, 2500 to

2800 ft/sec with increasing PLA. _et varied from Vjet ,,1800 to 2200 ft/sec. Figure 6(b) shows the relationship of

PLA to exhaust Mach numbers M 9 and Mje t. The values ofM9 varied between M9 - 1.71 to 1.74 and Mjet variedbetween Mje: = 1.08 to 1.30. The data for the exhaustMach numbers show a trend similar to the exhaust

velocities in Fig. 6(a). Additional data for the ground testsare listed in Table 4.

The data in the tables and figures are typical pointstaken from the many runs conducted in the study. Theymatched the desired altitudes and Mach numbers shown in

the flight matrix in Table 1. The data were selected fromtest points with stable engine conditions. These pointswere not averaged. The overall results show that theengine exhaust characteristics of interest for the acoustictest vary with M,_ altitude, and PLA. Tables 2, 3, and 4show that from the climb-to-cruise, ANOPP validation,

and ground tests the peak V9 values were approximately2800 to 2900 fl/sec.

Concluding Remarks

A series of acoustic tests were conducted, first, to

determine climb-to-cruise noise of aircraft with high noz-zle pressure ratios and second, to validate the AircraftNoise Prediction Program (ANOPP). An F-18 airplane,with the F404-GE-400 engine installed, was flown over arange of flight speeds and altitudes. From these tests, theengine data were analyzed to determine their exhaust char-acteristics. The flight tests produced a large engineexhaust characteristics database that was correlated withacoustic data and used to upgrade the ANOPP code. Thisnew database will aid in the design of fuutre high-speedcivil transport (HSCT) aircraft.

In summary, the climb-to-cruise test conditions atintermediate power produced engine exhaust conditionsthat varied from overexpanded to underexpanded. Thenozzle exit velocities ranged from approximately 2750 ft/sec up to a maximum of approximately 2800 ft/sec, andthen dropped to a value of approximately 2750 ft/sec. Thenozzle exit Mach numbers ranged between Mach 1.69 upto a maximum of Mach 1.81 and then dropped to a valueof Math 1.70. The maximum nozzle pressure ratio was5.24. For the ANOPP validation test points, the exhaustconditions were overexpanded and nozzle exit velocityranged from approximately 2550 to 2900 ft/sec. Nozzleexit Mach numbers ranged from approximately Mach 1.70to Mach 1.81. For the ground test conditions, nozzle exitvelocity varied from 2500 to 2800 ft/sec. Nozzle exitMnch number remained fairly constant at Math 1.70 overthe range of power levels tested. For the three tests:climb-to-cruise, ANOPP, and the ground run, at intermedi-ate power, maximum nozzle exit velocities were approxi-mately 2800 to 2900 ft/sec and nozzle exit Mach numberwas approximately Mach 1.70.

References

lUnited States, Federal Aviation Administration,

Code of Federal Regulations--Aeronautics and Space,vol. 14, pts. 1 to 59, Washington, DC, rev. Jan. 1, 1992.

2Vachal, John D., "High-Speed Civil Transport

Research and Technology Needs," SAE 901925, SAE

Aerospace Technology Conference and Exposition, Oct.

1990, pp. 1824-1831.

3Whitehead, Allen H., Jr., "Overview of Airframe

Technology in the NASA High-Speed Research Program,"

AIAA-91-3100, Sept. 1991.

4Kelly, J.J. and M.R. Wilson, "Signal Processing of

Jet Noise From Flyover Test Data," AIAA-93-0736, Jan.1993.

5powel, S.E, IV, "On the Leading Edge: Combining

Maturity and Advanced Technology on the 1=404 Turbofan

Engine," ASME, Transactions, J. of Engineering for GasTurbines andPower, vol. 113, Jan. 1991, pp. 1-10.

6Walton, James T. and Frank W. Burcham, Jr.,

Exhaust-Gas Pressure and Temperature Survey of F404-

GE-400 Turbofan Engine, NASA TM-88273, 1986.

7Preisser, J.S., R.A. Golub, J.M. Seiner, and C.A.

Powell, "Supersonic Jet Noise: Its Generation, Prediction,

and Effects on People and Structures," SAE 1990 Transac-

tions, J. of Aerospace, sec. 1, vol. 99, pt. 2, 1990,

pp. 1833-1847. (Also available as SAE 901927, 1990.)

SF404 In-Flight Thrust Calculation Deck, Program

no. 83112, General Electric Co., Aug. 1983.

C_.L..I

_ 00 oC o_ O0 o0

b_

i

.5

o,_

oral

c_

C_

0_ _

o

o0

8

°o°oo

oo°°,

°,°°o

6

0

0

_8

Q

I',- _ Ox

• ° ° °

• •

• ° _ °

m

.I

0

7

_8

I I I I

I I I I

I I I I

Fig. 1 F-18 aircraft powered by two F404-GE-400 engines.

Typical grounddata trajectory -_

North

NASADryden

Radar site •

920747

Fig. 2 Ground-wacking and array layout at Rogers Dry Lake, Edwards, California.

8

Nozzle exit Mach number (Mg)

Nozzle exit velocity (V9)

__ Fully expanded jet Mach number M.

Nozzle exit plane-_ Fldeal fully expanded no_le (Jet)

Direction of flight Fully expanded jet velocity (Vjet)

C_ Jet ___

flow • • • • ,rI _ _-. ,'_,,._:-_._ Shock ure

Nozzle throat J

B:_:dkbac_dnoise

930107

Fig. 3 Noise sources for an engine operating at high NPRs.

3000

2900

2800

2700

Exhaustvelocity, 2600

flJsec

2500

2400

23O0

2200.2

D

0"° I 12,300ft _ -__

I !o22,3_0 ft

I I I I I I I.3 .4 .5 .6 .7 .8 .9

Free-stream Mach number

(a) Nozzle exit and fully expanded jet velocity.

"'O'" Nozzle exit

veloc.y(vg)Fully expanded

Jetvelocity(vJet)

I1.0

930108

ExhaustMach

number

1.9

1.8

1.7

1.6

1o5 B

1.4-

1.3.2

"'O'" Nozzle exit

Mach number (Mg)

oo.oooOC_---_.. _ Fully expanded Jet

o'"'""'"'"" iI=,J__S Mach number (Mjet)

3800 ft 7300 ft

-r"II I I I I I I I

.3 .4 .5 .6 .7 .8 .9 1.0Free-stream Mach number

830109

(b) Nozzle exit and fully expanded jet Math number.

Fig. 4 Effect of Mach number and altitude on exhaust characteristics for climb-to-cruise test points, PLA setting atintermediate.

I0

3000 --

2800

2600

Exhaustvelocity, 2400

ft/sec

220O

2O00

1800.3

.,..,...0 .......

-o ................

-o

I I I I I I.4 .5 .6 .7 .8 .9

Free-stream Mach number930110

--0-- Nozzle exitvelocity (Vg)

Fully expanded

Jetveloclty (VJet)(PLA) Power lever angle

I1.0


2.0

1.8 -- o .......... (_ ....... 0

O-.....................0 -°°°

1.6 (102°)

IMachnumber --O-- Nozzle exit

1"4 F _ Mach number (M9)Fully expanded Jet

Mach number (Mjet)

1.2 F[_._._ _'_ (PLA) Power lever angle

/1.o/ I I I I I I I

•3 .4 .5 .6 .7 .8 .9 1.0Free-stream Mach number o3o111

Co) Nozzle exit and fully expanded jet Mach number.

Fig. 5 Effect of Mach number on F404 engine exhaust characteristics for ANOPP test points, PLA setting at power for

level flight (as noted).

11

Exhaustvelocity,

It/see

28O0

26OO

24O0

2200--

2000 --

18007O

• power

80 90 100 110

Power lever angle ,soln


Nozzle exit

velocity (3/9)Fully expanded

Jetvelocity (Vjet)

ExhaustMach

number

1.8

1.6 m

1.4--

1.2m

1.070

0 ........... 0 ....... 0 .................. 0

Nozzle exit

Uach number (M9)Fully expanded

JetMach number (Mjet)

i i8O 9O

Power lever angle

-'3

Intermediate power

I100

i110

930113

Co) Nozzle exit and fully expanded jet Math number.

Fig. 6 Effect of PLA on F404 engine exhaust characteristics for ground static test points, 2300-ft altitude.

12

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October 1993 Technical Memorandum

4. TrrLtANDSUBTrrLE 5. FUNDINGNUMBERS

Right-Determined Engine Exhaust Characteristics of an F404 Engine in an

F- 18 Airplane

6. AUTHOR(S)


7,PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(B)

NASA Dryden Flight Research FacilityP.O. Box 273

Edwards, California 93523-0273

9. SPONSORING/MONOTORING AGENCY NAME(S) AND ADORESS(ES)

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11.SUPPLEMENTARYNOTES

Prepared as AIAA-93-2543 for presentation at the AIAA/SAFJASME/ASEE 29th Joint Propulsion Conferenceand Exhibit, Monterey, California, June 28-30, 1993.

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13. ABSTRACT (Maximum 200 words)

Personnel at the NASA Langley Research Center (NASA-Langley) and the NASA Dryden Flight Research Fa-

cility (NASA-Dryden) have recently completed a joint acoustic flight test program. Several types of aircraft with

high nozzle pressure ratio engines were flown to satisfy a twofold objective. First, assessments were made of sub-

sonic climb-to-cruise noise from flights conducted at varying altitudes in a Mach 0.30 to 0.90 range. Second, us-

ing data from flights conducted at constant altitude in a Math 0.30 to 0.95 range, engineers obtained a high-

quality noise database. This database was desired to validate the Aircraft Noise Prediction Program and other sys-

tem noise prediction codes. NASA-Dryden personnel analyzed the engine data from several aircraft that were

flown in the test program to determine the exhaust characteristics. The analysis of the exhaust characteristics from

the F-18 aircraft will be reported in this paper. This paper presents an overview of the flight test planning, instru-

mentation, test procedures, data analysis, engine modeling codes, and results.

14.SUBJECTTERMS

Acoustics, Climb to cruise, Engine exhaust characteristics, Environmental impact

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